1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine blade with a cooling circuit.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, especially in an industrial gas turbine engine, a turbine section includes multiple stages of stator or guide vanes and rotor blades to extract mechanical energy from a hot gas flow passing through the turbine. Increasing the turbine inlet temperature can increase the turbine efficiency, and therefore the engine efficiency. However, the maximum turbine inlet temperature is limited to the material characteristics of the turbine airfoils, especially the first stage guide vanes and rotor blades, since these airfoils are exposed to the highest temperature.
In order to allow for a higher gas flow temperature, the turbine airfoils include complex internal cooling circuits to provide the maximum amount of cooling for the airfoil while making use of the minimum amount of cooling air in order to maximize the efficiency of the turbine and therefore the engine. In a prior art turbine blade with near wall cooling, the airfoil main body includes radial flow channel plus re-supply holes in conjunction with film discharge cooling holes from the near wall channel. In this prior art airfoil, spanwise (the direction from root to tip) and chord wise (the direction from leading edge to trailing edge) cooling flow control due to airfoil external hot gas temperature and pressure variation is difficult to achieve. In addition, a single radial channel flow is not the best method of utilizing cooling air since it results in low convective cooling.
It is therefore an object of the present invention to provide for a turbine airfoil with a cooling circuit that will reduce the main body metal temperature and therefore reduce the cooling flow requirement and improve the turbine efficiency.